1. Field of the Invention
This invention resides in the technology of nozzle design for rocket propulsion systems.
2. Description of the Prior Art
Rocket-powered launch vehicles require high thrust at takeoff due to the large amount of unburned fuel initially present in the vehicle. For vehicles that are launched from the earth's surface, takeoff typically occurs at sea level while the vehicle performs its primary mission at high altitude where the external pressure is lower and is often at high vacuum. To perform its primary mission effectively, the vehicle must produce a high specific impulse (Isp), i.e., a high ratio of thrust to the weight of fuel consumed in a unit of time. This is most readily achieved when the engine has a nozzle with a high area ratio, which is the ratio of the area at the nozzle exit to the area at the throat. Nozzles with high area ratios tend to produce relatively low thrust at sea level, however, because of a reverse pressure differential near the nozzle exit that occurs when the wall pressure is below ambient pressure. This reverse pressure differential produces a negative thrust component in the portion of the nozzle near the exit, i.e., a thrust component whose direction is opposite to the forward direction of the vehicle. This negative component reduces the total thrust produced by the nozzle.
One method in the prior art to eliminate this negative component of the sea level thrust without compromising the thrust in a high vacuum environment is the use of a nozzle of variable area, i.e., one in which the area at the exit is reduced for launch and then gradually increased during ascent. The variation is achieved by constructing the nozzle with the capability of adjustments to the contour, area ratio and length of the nozzle as the vehicle altitude increases. Features such as these add considerable complexity and weight to the engine construction, however, and they are less than fully successful since the nozzle in most cases continues to produce less thrust at sea level than at vacuum. Other methods have included the use of combination-type engines using different fuels at different stages. Typical such combinations are kerosene-fueled engines combined with engines derived from the Space Shuttle Main Engine (SSME), kerosene-fueled engines combined with hydrogen-fueled engines such as the Russian RD-701 engine, the dual-fuel-dual-expander engine concept described by Beichel, R., in U.S. Pat. No. 4,220,001 (issued Sep. 2, 1980), and the dual-thrust rocket motor of Bornstein, L., U.S. Pat. No. 4,137,286 (issued Jan. 30, 1979) and U.S. Pat. No. 4,223,606 (issued Sep. 23, 1980). The Beichel engine requires a complex nozzle design that incorporates two thrust chambers, while the Bornstein motor achieves dual thrust by using separate sustainer and booster propellant grains in the combustion chamber, together with an igniter and squib that are inserted into the grain itself. A further alternative is the introduction of secondary combustion gas near the wall of the divergent section of the nozzle, as described by Bulman, M., in U.S. Pat. No. 6,568,171 (issued May 27, 2003).
Of further possible relevance to this invention are space vehicles, notably those that are designed to undergo ascent, descent, or both in high-vacuum environments such as the surface of the moon, either to return to earth or to enter a lunar orbit. These vehicles require very deep throttling upon approaching the landing surface and the need to vary thrust upon takeoff from a very high level at the take-off surface to a lower level when landing on the moon.